Review on active thermal protection and its heat transfer for airbreathing hypersonic vehicles

The current work is an experimental investigation of the dependence of film cooling effectiveness on the injection Mach number, velocity, and mass flux. The freestream Mach number is 2.4, and the injection Mach numbers range from 1.2 to 2.2 for both air and helium injection. The adiabatic wall temperature is measured directly. The injection velocity and mass flux are varied by changing the total temperature and Mach number while maintaining matched pressure conditions between the injected flow and that of the freestream. The total temperature of the freestream is 295 K, and for the injection it ranges from 215-390 K. The results indicate an increase in film cooling effectiveness as the injection rate is increased. With the exception of heated helium runs, larger injection Mach numbers slightly increase the effective cooling length per mass injection rate. The results for helium injection indicate an increase in effectiveness as compared to that for air injection. Heated injection, with the injectant to freestream velocity ratios greater than 1, exhibit a rise in wall temperature downstream of the slot resulting in effectiveness values greater than 1. The experimental results are also compared with earlier studies in the literature.

Influence of injectant Mach number and temperature on supersonic film cooling

Research progress on active thermal protection for hypersonic vehicles

Experimental investigation on convective heat transfer from a horizontal miniature tube to methane at supercritical pressures

Large-Eddy Simulation of Shock/Cooling-Film Interaction

The supersonic film cooling was tested in the Mach 2.35 wind tunnel to investigate the effect of the external shock wave on the film cooling. The coolant was injected with sonic speed. The weak shock wave with the pressure ratio of 1.21 did not reduce the film cooling effectiveness. The stronger shock wave with the pressure ratio of 1.44 decreased the effectiveness of the film cooling in the restricted region. The decrease of the effectiveness was mainly the result of the increase of the adiabatic wall temperature by the decrease of the local Mach number. The increase of the heat transfer coefficient must be considered as well as that of the adiabatic wall temperature. In the region of the interaction, energy and mass were not transferred, but the momentum was transferred from the primary flow to the coolant.

Experimental studies of supersonic film cooling with shock wave interaction

Experimental Studies of Supersonic Film Cooling with Shock Wave Interaction (II)

LOX/methane staged combustion rocket combustor investigation

The required operating conditions for the present rocket thrust chamber for rocketramjet combined-cycle engine are 1) chamber pressure Pc = 5.0 MPa, mixture ratio O/F = 7 for the ejector-jet mode, the scramjet mode, and the rocket mode, 2) Pc = 0.6 MPa, O/F = 3 for the ramjet mode. Stable operation of a rocket engine in a wide range of both Pc and O/F is required in order to adjust the operation mode in accordance with flight speed. Gaseous hydrogen and gaseous oxygen were used as the propellant. In the previous study, the hydrogen flow rate during firing tests decreased due to the thermal deformation of the faceplate. In this study, a new designed injector which has eight hydrogen injection holes arranged around an oxygen post was proposed for the rocket thrust chamber of the rocketramjet combined-cycle engine. The hydrogen flow rate during firing tests was almost the same as that during the cold flow tests. Performances of 0.84 to 0.88 in C* efficiency was achieved under the condition of O/F = 6.5 to 7.5 using the thrust chamber of L* = 0.33 m. However performances of about 1.0 in C* efficiency were achieved under the conditions of O/F = 4 to 6.5 using the thrust chamber of L* = 0.99 m.

Injectors and Combustion Performance of Rocket Thruster for Rocket-Ramjet Combined-Cycle Engine Model

A modified sidewall-compression-type scramjet engine was tested under M8 flight condition. A strut with a long tail led to a 60% higher maximum thrust increment than a strut with a truncated tail, while allowable fuel flow rate (i.e., the flow rate at engine unstart limit) was almost identical. To increase the allowable fuel flow rate and the maximum thrust increment due to combustion without engine unstart, distributed injection in vertical direction and installation of ramps in the isolator between the inlet and the combustor were applied. Distributed fuel injection led to a 30% higher maximum thrust increment than the uniform fuel injection, and the installation of the ramps also gave a 30% higher one than that in the engine without ramp. Gas-sampling at the exit of the engine with distributed fuel injection showed 88% of bulk combustion efficiency. Specific impulses of the maximum thrust increments, δIsp , of the modified engine were 60–80% higher than that of the original engine.

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Fumiei Ono

Papers

Experimental studies of supersonic film cooling with shock wave interaction

Experimental Studies of Supersonic Film Cooling with Shock Wave Interaction (II)

LOX/methane staged combustion rocket combustor investigation

Injectors and Combustion Performance of Rocket Thruster for Rocket-Ramjet Combined-Cycle Engine Model

Modified Water-Cooled Scramjet Engine Tested under M8 Flight Condition.